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| United States Patent |
5,087,305
|
|
Chang
|
February 11, 1992
|
Fatigue crack resistant nickel base superalloy
Abstract
An alloy is disclosed which has been found to lend itself particularly well
to thermomechanical processing. The alloy is strengthened by precipitates
similar to those of Inconel 718 but the alloy matrix of the composition is
a nickel-chromium-cobalt matrix rather than the nickel-chromium-iron
matrix of the Inconel alloy. Also the alloy has grains of average diameter
of 35 .mu.m or larger. The fatigue resistance, tensile strength and the
rupture strength of the alloy is improved to a remarkable degree as a
result of the thermomechanical processing. The thermomechanical processing
is carried out below the recrystallization temperature to prevent
nucleation of fine grains. The residual strains from the thermomechanical
processing or cold working provides the remarkably favorable combination
of alloy properties which are found.
| Inventors:
|
Chang; Keh-Minn (Schenectady, NY)
|
| Assignee:
|
General Electric Company (Schenectady, NY)
|
| Appl. No.:
|
215189 |
| Filed:
|
July 5, 1988 |
| Current U.S. Class: |
148/410; 148/677 |
| Intern'l Class: |
C22C 019/05 |
| Field of Search: |
148/410,12.7 N,428
420/448
|
References Cited
U.S. Patent Documents
| 3046108 | Jul., 1962 | Eiselstein | 420/448.
|
| 3372068 | May., 1968 | White et al. | 148/162.
|
| 4140555 | Feb., 1979 | Garcia et al. | 148/32.
|
| Foreign Patent Documents |
| 260510 | Mar., 1988 | EP.
| |
| 2133186 | Feb., 1972 | DE.
| |
| 3427206 | Jul., 1985 | DE.
| |
| 2223470 | Mar., 1974 | FR.
| |
Primary Examiner: Dean; R.
Attorney, Agent or Firm: Rochford; Paul E., Davis, Jr.; James C., Magee, Jr.; James
Claims
What is claimed and sought to be protected by Letters Patent of the United
States is as follows:
1. A structural article having high strength and low fatigue crack
propagation rate which comprises
an article formed of a composition consisting essentially of the following
in parts by weight:
______________________________________
Concentration
Ingredient From About To About
______________________________________
Nickel balance
Chromium 16 22
Cobalt 8 14
Molybdenum 2.0 4.0
Aluminum 0.2 0.9
Titanium 0.5 1.5
Tantalum 3.5 4.5
Niobium 3.5 4.5
Carbon 0.0 0.05
Boron 0.002 0.015
______________________________________
the composition having been recrystallized and aged and having grains of
minimum average diameter of about 35 microns, and
the grains of the article being deformed by a mechanical working to change
the shape of the article by at least 15%.
2. The article of claim 1 in which the change in shape is at least 20%.
3. The article of claim 1 in which the change in the shape is at least 25%.
4. The article of claim 1 in which the change of shape is at least 35%.
5. A structural article having high strength and low fatigue crack
propagation rate which comprises
an article formed of a composition consisting essentially of the following
in parts by weight:
______________________________________
Ingredient Concentration
______________________________________
Nickel balance
Chromium 12
Cobalt 18
Molybdenum 3
Aluminum 0.5
Titanium 1
Tantalum 4
Niobium 4
Carbon 0.015
Boron 0.01.
______________________________________
the composition having been recrystallized and aged and having grains of
minimum average diameter of about 35 microns, and
the grains of the article being deformed by a mechanical working to change
the shape of the article by at least 15%.
Description
RELATED APPLICATIONS
The subject application relates generally to the subject matter of
applications Ser. Nos. 907,275 and 907,550, filed concurrently on Sept.
15, 1986 which applications are assigned to the same assignee as the
subject application herein.
The texts of these related applications is incorporated herein by
reference.
BACKGROUND OF THE INVENTION
It is well known that nickel based superalloys are extensively employed in
high performance environments. Such alloys have been used extensively in
jet engines and in gas turbines where they must retain high strength and
other desirable physical properties at elevated temperatures of a
1000.degree. F. or more.
The strength of these alloys is related to the presence of a strengthening
precipitate, which in many cases is a .delta.' precipitate or .delta."
precipitate. More detailed characteristics of the phase chemistry of
precipitates are given in "Phase Chemistries in
Precipitation-Strengthening Super-alloy" by E. L. Hall, Y. M. Kouh, and K.
M. Chang [Proceedings of 41st. Annual Meeting of Electron Microscopy
Society of America, August 1983 (p. 248)].
The following U.S. patents disclose various nickel-base alloy compositions,
some of which contain such precipitates: U.S. Pat. No. 2,570,193; U.S.
Pat. No. 2,621,122; U.S. Pat. No. 3,046,108; U.S. Pat. No. 3,061,426; U.S.
Pat. No. 3,151,981; U.S. Pat. No. 3,166,412; U.S. Pat. No. 3,322,534; U.S.
Pat. No. 3,343,950; U.S. Pat. No. 3,575,734; U.S. Pat. No. 3,576,681; U.S.
Pat. No. 4,207,098 and U.S. Pat. No. 4,336,312. The aforementioned patents
are representative of the many alloying situations reported to date in
which many of the same elements are combined to achieve distinctly
different functional relationships between the elements such that phases
form which provide the alloy system with different physical and mechanical
characteristics. Nevertheless, despite the large amount of data available
concerning the nickel-base alloys, it is still not possible for workers in
the art to predict with any degree of accuracy the physical and mechanical
properties that will be displayed by certain concentrations of known
elements used in combination to form such alloys even though such
combination may fall within broad, generalized teachings in the art,
particularly when the alloys are processed using heat treatments different
from those previously employed.
A significant development in the alloys for use at high temperature was
made in 1962 with the development of the IN718 alloy by H. L. Eiselstein
at the International Nickel Company. The Eiselstein patent U.S. Pat. No.
3,046,108 resulted from this discovery and was the basis for the
commercial production of the alloy IN718 which is still produced and used
very extensively commercially. This alloy was characterized by the
presence therein of a substantial quantity of .delta." precipitate.
Studies of the alloy and of the precipitate are contained in the following
papers:
"Alloy 718: The Workhorse of Superalloys", by Robert R. Irving, Iron Age,
June 10, 1981;
"Metallurgy of a Columbium-Hardened Nickel-Chromium-Iron Alloy", by
Eiselstein, Advances in the Technology of Stainless Steels, pp. 62-79;
"Identification of the Strengthening Phase in "Inconel" Alloy 718" by
Kotval, Transactions of the Metallurgical Society of AIME, Vol. 242,
August 1968, pp. 1764-65;
"Precipitation of Nickel-Base Alloy 718", by Paulonis et al., Transactions
of the ASM, Vol. 62, 1969, pp. 611-622"
"Effect of Grain Boundary Denudation of Gamma Prime on Notch-Rupture
Ductility of Inconel Nickel-Chromium Alloys X-750 and 718", by E. L.
Raymond, Transactions of the Metallurgical Society of AIME, Vol. 239,
Sept. 1967, pp. 1415-1422.
Essentially, no improvements were made in the alloy for approximately 25
years from the date when the Eiselstein application was filed on the IN718
alloy in November, 1958. Recently, however, a unique and unusual
improvement was made in alloys which are strengthened by .delta."
precipitate and the description of this new class of alloys resulting from
the discovery is described in the UK Patent Application GB2148323A.
It is known that some of the most demanding sets of properties for
superalloys are those which are needed in connection with jet engine
construction. Of the sets of properties which are needed those which are
needed for the moving parts of the engine are usually greater than those
needed for static parts although the sets of needed properties are
different for the different components of an engine.
Because some sets of properties have not been attainable in cast alloy
materials, resort is sometimes had to the preparation of parts by powder
metallurgy techniques. However, one of the limitations which attends the
use of powder metallurgy techniques in preparing moving parts for jet
engines is that of the purity of the powder. If the powder contains
impurities such as a speck of ceramic or oxide the place where that speck
occurs in the moving part becomes a latent weak spot where a crack may
initiate or it becomes a latent crack.
To avoid problems with impure powder and similar problems it is sometimes
preferred to form moving parts of jet engines such as disks with alloys
which can be cast and wrought.
A problem which has been recognized to a greater and greater degree with
many such nickel based superalloys is that they are subject to formation
of cracks or incipient cracks, either in fabrication or in use, and that
the cracks can actually initiate or propagate or grow while under stress
as during use of the alloys in such structures as gas turbines and jet
engines. The propagation or enlargement of cracks can lead to part
fracture or other failure. The consequence of the failure of the moving
mechanical part due to crack formation and propagation is well understood.
In jet engines it can be particularly hazardous.
However, what has been poorly understood until recent studies were
conducted was that the formation and the propagation of cracks in
structures formed of superalloys is not a monolithic phenomena in which
all cracks are formed and propagated by the same mechanism and at the same
rate and according to the same parameters and criteria. By contrast the
complexity of the crack generation and propagation and of the crack
phenomena generally, and the interdependence of such propagation with the
manner in which stress is applied, is a subject on which important new
information has been gathered in recent years. The period during which
stress is applied to a member to develop or propagate a crack, the
intensity of the stress applied, the rate of application and of removal of
stress to and from the member and the schedule of the application was not
well understood in the industry until a study was conducted under contract
to the National Aeronautics and Space Administration. This study is
reported to a technical report identified as NASA CR-165123 issued from
the National Aeronautics and Space Administration in August 1980,
identified as "Evaluation of the Cyclic Behavior of Aircraft Turbine Disk
Alloys", Part II, Final Report, by B. A. Cowles, J. R. Warren and F. K.
Hauke, and prepared for the National Aeronautics and Space Administration,
NASA Lewis Research Center, Contract NAS3-21379.
A principal unique finding of the NASA sponsored study was that the rate of
propagation based on fatigue phenomena or in other words the rate of
fatigue crack propagation (FCP) was not uniform for all stresses applied
nor to all manners of applications of stress. More importantly, the
finding was that fatigue crack propagation actually varied with the
frequency of the application of stress to the member where the stress was
applied in a manner to enlarge the crack. More surprising still, was the
finding from the NASA sponsored study that the application of stress of
lower frequencies rather than at the higher frequencies previously
employed in studies, actually increased the rate of crack propagation. In
other words the NASA study revealed that there was a time dependence in
fatigue crack propagation. Further the time dependence of fatigue crack
propagation was found to depend not on frequency alone but on the time
during which the member was held under stress or a so-called hold-time.
Following the discovery of this unusual and unexpected phenomena of
increased fatigue crack propagation at lower stress frequencies there was
some belief in the industry that this newly discovered phenomena
represented an ultimate limitation on the ability of the nickel based
superalloys to be employed in the stress bearing parts of the turbines and
aircraft engines and that all design effort had to be made to design
around this problem.
However, it has been discovered that it is feasible to construct parts of
nickel based superalloys for use at high stress in turbines and aircraft
engines with greatly reduced crack propagation rates.
The development of the superalloy compositions and methods of their
processing of this invention focuses on the fatigue property and addresses
in particular the time dependence of crack growth.
Crack growth, i.e., the crack propagation rate, in high-strength alloy
bodies is known to depend upon the applied stress (.sigma.) as well as the
crack length (a). These two factors are combined by fracture mechanics to
form one single crack growth driving force; namely, stress intensity K,
which is proportional to .sigma..sqroot.a. Under the fatigue condition,
the stress intensity in a fatigue cycle represents the maximum variation
of cyclic stress intensity (.DELTA.K), i.e., the difference between
K.sub.max and K.sub.min. At moderate temperatures, crack growth is
determined primarily by the cyclic stress intensity (.DELTA.K) until the
static fracture toughness K.sub.IC is reached. Crack growth rate is
expressed mathematically as da/dN .varies.(.DELTA.K).sup.n. N represents
the number of cycles and n is a constant which is between 2 and 4. The
cyclic frequency and the shape of the waveform are the important
parameters determining the crack growth rate. For a given cyclic stress
intensity, a slower cyclic frequency can result in a faster crack growth
rate. This undesirable time-dependent behavior of fatigue crack
propagation can occur in most existing high strength superalloys.
According to this hold time pattern, the stress is held for a designated
hold time each time the stress reaches a maximum in following the normal
sine curve. This hold time pattern of application of stress is a separate
criteria for studying crack growth. This type of hold time pattern was
used in the NASA study referred to above.
The design objective is to make the value of da/dN as small and as free of
time-dependency as possible.
It is pointed out in copending application Ser. No. 907,550, filed Sept.
15, 1986 that time dependent fatigue crack propagation can be reduced
significantly by a thermal treatment of .delta.' strengthened nickel base
superalloys which have more than 35 volume percent of strengthening
precipitate. As is pointed out in this copending application, the method
involves a high temperature solutioning (supersolvus) solutioning of the
.delta.' precipitate followed by a controlled cooling at less than
250.degree. F. per minute.
However, it has been found that the method of copending application Ser.
No. 907,550 does not yield the beneficial results taught in that
application when the method is applied to alloys with low precipitate
content. For example, the method does not produce the fatigue crack
propagation reduction when applied to Waspalloy or to IN718 alloy.
Waspalloy is .delta.' hardened and has less than 35 volume percent and
preferably about 30 volume percent .delta.' precipitate. IN718 is mainly
.delta.'' hardened and has less than 35 volume percent and preferably
about 20 percent by volume of .delta.' precipitate.
I have done extensive studies on alloys of such lower .delta.' or .delta."
precipitate content and have heat treated these alloys according to a
variety of schedules which restrict fatigue crack propagation properties
of alloys having higher precipitate content but without significant
beneficial effect. I have found that none of these heat treatments develop
different or advantageous microstructures or result in any significant
reduction in fatigue crack propagation.
A second copending application Ser. No. 907,275, also filed Sept. 15, 1986,
discloses a method for processing a superalloy containing a lower
concentration of strengthening precipitate. The method of this copending
application produces materials with a superior set or combination of
properties for use in advanced engine disc applications. Properties which
are conventionally needed for materials used in disc applications include
high tensile strength and high stress rupture strength. These properties
are achieved in the practice of the method of the copending application
Ser. No. 907,275 and, in addition, the alloy prepared by the methods of
the copending application exhibit a desirable property of resisting crack
growth propagation. Such ability to resist crack growth in essential for
the component low cycle fatigue life or LCF. In addition to this superior
set of properties as outlined above, the alloy processed according to the
method of the Ser. No. 907,275 copending application displays good
forgeability and such forgeability permits greater flexibility in the use
of various manufacturing processes needed in formation of parts such as
discs for jet engines. Superalloys with lower ranges of precipitate
content generally have good forgeability and can be subjected to
thermomechanical processing. The differences in the results obtained by
certain thermomechanical processings on mechanical properties, like
strength and rupture life, are known to a degree. However, prior to the
teaching of the copending application Ser. No. 907,275 nothing was known
of the influence if any of thermomechanical processing on time-dependent
fatigue crack propagation or the rates of such propagation.
As alloy products for use in turbines and jet engines have developed it has
become apparent that different sets of properties are needed for parts
which are employed in different parts of the engine or turbine. For jet
engines, the material requirements of more advanced aircraft engines
continue to become more strict as the performance requirement of the
aircraft engines are increased. The different requirements are evidenced,
for example, by the fact that many blade alloys display very good high
temperature properties in the cast form. However, the direct conversion of
cast blade alloys into disc alloys is very unlikely because blade alloys
display inadequate strength at intermediate temperatures of about
700.degree. C. Further, the blade alloys have been found very difficult to
forge and forging has been found desirable in the fabrication of blades
from disc alloys. Moreover, the crack growth resistance of disc alloys has
not been evaluated.
Accordingly, to achieve increased engine efficiency and greater
performance, constant demands are made for improvements in the strength
and temperature capabilities of disc alloys as a special group of alloys
for use in aircraft engines. Now, these capabilities must be coupled with
low fatigue crack propagation rates and a low order of time dependency of
such rates.
While the copending application Ser. No. 907,275 dealt with the
improvements which could be made in existing alloys of low precipitate
concentration through the thermomechanical processing, there was no
disclosure of any alloy in the copending application which was
particularly adapted to be benefitted by the application of the
thermomechanical processing of the copending application or of novel
results of the application of such processing to an alloy so adapted.
The present invention provides a alloy which is particularly adapted and
suited to the processing by thermomechanical treatment taught in the
copending application to achieve a unique and remarkable combination and
set of properties.
BRIEF DESCRIPTION OF THE INVENTION
It is accordingly one object of the present invention to provide
nickel-base superalloy products which are more resistant to cracking.
Another object is to provide novel alloy which is particularly suited to
increasing the high temperature capability thereof.
Another object is to provide articles for use under cyclic high stress
which are more resistant to rupture.
Another object is to provide a method for reducing the time dependency of
fatigue cracking in combination with unique alloys having higher strength.
Another object of the present invention to provide the combination of a
novel composition and method which permits the novel superalloys to
display increased strength and increased rupture properties.
Another object is to provide an alloy which has principally precipitate
strengtheners adapted to be processed into a condition in which the high
temperature capabilities of the alloy is emphasized.
Other objects will be in part apparent and in part pointed out in the
description which follows.
In one of its broader aspects, objects of the present invention can be
achieved by providing an alloy having a composition in weight percent
essentially as follows:
______________________________________
Concentration in
weight percent
Ingredient From about
To about
______________________________________
Nickel balance
Chromium 16 22
Cobalt 8 14
Molybdenum 2.0 4.0
Aluminum 0.2 0.9
Titanium 0.5 1.5
Tantalum 3.5 4.5
Niobium 3.5 4.5
Carbon 0.0 0.05
Boron 0.002 0.015
______________________________________
The alloy of the present invention is strengthened by precipitates similar
to those of Inconel 718. However, the alloy matrix of the composition is a
nickel-chromium-cobalt matrix rather than the nickel-chromium-iron matrix
of the Inconel 718 alloys.
By balance nickel as used herein it is meant that the balance is
predominantly nickel but that the composition may contain minor amounts of
other elements such as iron, magnesium and other elements as impurities or
as minor additives so long as the presence of the other elements does not
detract from or interfere with the beneficial properties of the alloy as
taught herein.
The alloy, which is set out above, has been found to be particularly suited
and adapted to receive the thermomechanical processing treatments as set
for the in copending application Ser. No. 907,275 which application is
incorporated herein by reference. The result of the development of this
designated composition and the application of the thermomechanical
processing is to achieve a composition with crack growth resistance that
has improved high temperature strength and temperature capability superior
to commercial alloys which have received the benefit of the
thermomechanical processing described in the copending U.S. Ser. No.
907,275 application.
It should be emphasized that the novelty of the subject invention resides
principally in the finding that this alloy, when coupled with the
thermomechanical processing of the copending application, yields unique
and novel properties. Novelty exists because the application of the same
thermomechanical processing to other alloys does not permit the
achievement of the superior strength and combination of other properties
developed in the subject alloy. In fact, there is no other alloy known to
the inventor which has the capacity for achieving the combination of
strength and other properties which the alloy of this invention can
achieve through the thermomechanical processing.
The sample is then given a solution heat treatment at a temperature above
the recrystallization temperature if the grain structure of the alloy is
smaller grains of average diameter of less than 35 .mu.m. The sample may
be aged following the solution heat treatment.
The sample must have acquired a recrystallized equiaxed grain structure
from the heat treatment and should have a strength which is essentially
normal for the alloy. The grain size should preferably be of the order of
35 .mu.m average diameter or larger.
The alloy sample is then subjected to mechanical working to distort the
grains thereof.
The mechanical working can be by a cold working as by a forging or by a
rolling or by a combination of cold working steps.
Alternatively, one or more steps of the working may be accompanied by a
heating at a temperature below the recrystallization temperature. The
heating is preferably of a type and to an extent which facilitates and
enhances the deformation of the grains of the alloy sample.
Any heating which results in a recrystallization or refinement of the grain
structure, should be avoided and, if it cannot be avoided entirely, then
it should be minimized.
However, the sample may be given an aging heat treatment which does not
result in recrystallization and which does not cancel the deformation of
the grains. The alloy can be fully hardened to develop its full strength
through aging treatment.
BRIEF DESCRIPTION OF THE DRAWINGS
In the description which follows clarity of understanding will be gained by
reference to the accompanying drawings in which:
FIGS. 1-7 are graphic (log-log plot) representations of fatigue crack
growth rates (da/dN) obtained at various stress intensities (.DELTA.K) for
different alloy compositions at elevated temperatures under cyclic stress
applications at a series of frequencies one of which cyclic stress
applications includes a hold time at maximum stress intensity.
FIG. 8 is a graph in which temperature in degrees F is allotted against
stress in ksi and displaying 100 hours rupture life values for alloys
given different thermomechanical processing treatments.
DETAILED DESCRIPTION OF THE INVENTION
In the copending application Ser. No. 907,275 it was brought out that it is
possible to impart to nickel-base superalloys having relatively lower
content of precipitate, desirable sets of properties, including low
fatigue crack propagation rates. It was found and disclosed in the
copending application that superalloys having lower concentrations of
precipitate of the order 35 volume percent or less can be treated by
thermomechanical processing to impart improvements to properties of the
alloys and specifically to the fatigue crack propagation rate for the
alloys.
However, this method was described as applied to existing alloys such as
the IN718 alloy. There was no disclosure of an alloy which was found to
have its properties particularly enhanced by thermomechanical processing.
The subject application teaches an alloy which has been found to have the
unique property of being particularly suited and adaptable to being
benefitted by the application of thermomechanical processing essentially
as taught in the copending application Ser. No. 907,275.
EXAMPLE 1
This example is essentially identical to Example 1 of U.S. Ser. No. 907,275
and deals with thermomechanical processing of a conventional alloy and
specifically IN718.
Several IN718 heats were prepared by conventional vacuum induction melting.
The melts were solidified and the ingots so formed were homogenized by
heating at 1200.degree. C. for 24 hours. The ingots were forged into
plates according to conventional practice for nickel base wrought
superalloys. The chemical composition of specific IN718 alloy employed in
these examples is set forth in Table I below:
TABLE I
______________________________________
Chemical Composition of Inconel 718
Element
wt. %
______________________________________
Ni bal.
Cr 19.0
Fe 18.0
Mo 3.0
Nb 5.1
Ti 0.9
Al 0.5
C 0.04
B 0.005
______________________________________
A metallographic study of the samples indicated that the IN718 alloy starts
to recrystallize when subjected to a temperature higher than 950.degree.
C.
The forged plates were subjected to standard heat treatment including a
solutioning at 975.degree. C. for one hour and a double aging at
720.degree. C. for eight hours. After the eight hour aging the samples
were furnace cooled at 620.degree. C. for an additional ten hours aging.
The material of the resulting forged plates was found to have a
recrystallized equiaxed grain structure of at least 35 .mu.m average
diameter. The strength of the forged samples was measured from room
temperature up to 700.degree. C. and was found to be similar in strength
to that of standard reference material.
Time dependent fatigue crack propagation was evaluated at 593.degree. C.
using three different fatigue waveforms similar to those used in the NASA
study. The first was a three second sinusoidal waveform and the second was
a 180 second sinusoidal waveform. The third was a 177 second hold at the
maximum load of three second sinusoidal cycle. The maximum to minimum load
ratio was set R=0.05 so that the maximum was 20.times.twenty fold higher
than the minimum load applied. Data was taken from the studies of the time
dependent fatigue crack propagation and the data is plotted in FIG. 1. The
results demonstrate and it can be observed from the plot that the crack
growth rate da/dN increases by a factor of six to eight times when the
fatigue cycle is changed from 3 seconds to 180 seconds. The hold time
cycle accelerates the crack growth rate by a factor of 20.
EXAMPLES 2 and 3
This example pertains to the application of the process of copending
application Ser. No. 907,725. to the commercially available alloy IN718 as
taught in the copending application.
Plates were prepared as described in Example 1 of alloy IN718. The plates
were prepared by vacuum induction melting followed by homogenization and
forging as described in the Examples above.
For Example 2 an alloy plate so prepared was cold rolled 20%. Test data was
taken of fatigue crack propagation rates for this 20% cold rolled sample
and the results are plotted in FIG. 2.
For Example 3 an alloy plate prepared as described above was cold rolled
through a 40% reduction in thickness. Fatigue crack propagation rate data
was taken for this sample and the data is plotted in FIG. 3.
It will be observed from examination and consideration of FIGS. 2 and 3
that there is a significant improvement in the fatigue crack propagation
time dependence. In other words the samples are found to be more
independent of time relationships of the testing at the three different
cycles and particularly at a 3 second cycle versus the 180 second cycle
versus the 3 second cycle with the 177 second hold period at maximum load.
The method of this example was described as applied to existing alloys and
specifically the IN718 alloy. There was no disclosure in copending
application Ser. No. 907,275 of the discovery of an alloy specifically
adapted to have its properties enhanced by thermomechanical processing.
The subject application teaches a alloy which has been discovered to have
the unique property of being particularly suited and adaptable to being
benefitted by the application of thermomechanical processing essentially
as taught in the copending application Ser. No. 907,275.
EXAMPLE 4
A sample of a different alloy was prepared for test. The procedures of
sample preparation are set out below. The composition prepared had the
composition as set forth in Table II.
TABLE II
______________________________________
Ingredient Nominal-CH84 Composition in wt %
______________________________________
Nickel balance
Chromium 12.00
Cobalt 18.00
Molybdenum 3.00
Aluminum 0.50
Titanium --
Tantalum --
Niobium 5.00
Carbon 0.015
Boron 0.01
______________________________________
The composition is described as nominal in that the ingredients were added
to achieve the percentages which are listed in Table II. The composition
was prepared by conventional vacuum induction melting. The melts were
solidified and the ingots so formed were homogenized by heating at
1200.degree. C. for 24 hours. The ingots were forged into plates according
to conventional practice for nickel-base wrought superalloys.
The samples were then subjected to the thermomechanical processing as
described in the copending application Ser. No. 907,275. In order to
simplify the thermomechanical processing, the forged plates were subjected
to different degrees of cold rolling. A 15% reduction by cold rolling was
designated D. A 25% reduction by cold rolling was designated E and a 35%
reduction in thickness by cold rolling was designated F.
Subsequent age treatments of 725.degree. C. for 8 hours and a furnace
cooling to 650.degree. C. and heating for 10 hours at that temperature
were applied to samples directly after the rolling.
The samples which were rolled to impart the three different degrees of
reduction were then tested for fatigue crack growth rate. The fatigue
crack growth rate was measured at 1100.degree. F. by using three fatigue
wave forms. A first being a 3 second sinusoidal cycle; a second being a
180 second sinusoidal cycle; the third being a 177 second hold cycle at
the maximum load of a 3 second cycle. This fatigue crack growth rate
measurements were essentially the same as those conducted in the copending
application Ser. No. 907,275 and in Example 1 above.
The results of the fatigue crack growth rate measurements for the sample D
given the 15% cold roll reduction and the sample E given the 25% cold roll
reduction are plotted in FIGS. 4 and 5. It is evident from FIGS. 4 and 5
that there was much less scatter of the test results based on the
differences in the test cycle applied than there was for the test samples
of Example 1 as these test results were plotted in FIG. 1. The reduction
in scatter is similar to that found in the FIGS. 2 and 3 developed from
the cold rolling reduction of the IN718 alloy specimen of Examples 2 and 3
above.
EXAMPLE 5
A heat was prepared to contain the composition as set forth in Table III
below in parts by weight.
TABLE III
______________________________________
Ingredient CH83 Composition in wt %
______________________________________
Nickel balance
Chromium 12.00
Cobal 18.00
Molybdenum 3.00
Aluminum 0.50
Titanium 1.00
Tantalum 4.00
Niobium 4.00
Carbon 0.015
Boron 0.01
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This composition contained the titanium and tantalum which where absent
from the composition of Example 4 above. This composition is within the
scope of the compositions taught in U.K. Patent Application GB2144323A.
The heat was processed through the preparation and thermal processing
procedures as described in Example 1 above. The grains of the
recrystallized alloy should preferably be at least 35 .mu.m in average
diameter.
Samples of the material were then subjected to thermomechanical processing
as also described in Example 2 above. Again a sample given a 15% reduction
by cold rolling was designated D. A 25% reduction by cold rolling was
designated F and a 35% reduction in thickness by cold rolling areas
designated F.
Samples of these thermomechanically processed alloys were subjected to
fatigue crack propagation testing as described in Examples 1 and 2 and the
results of the tests are plotted in FIGS. 6 and 7 for samples E and F. As
will be evident from a study of the results plotted in FIGS. 6 and 7 there
is very little time dependence of the fatigue crack propagation and
accordingly very little scatter of the data points of the plot, and
particularly of the data of FIG. 7 for the 35% cold rolled sample 83F.
EXAMPLE 6
The high temperature tensile properties of the alloys CH84 of Example 4 and
CH83 of Example 5 were measured and the results are given in Table IV.
Also in Table IV there is a listing of data which was obtained from
measurements on samples of Inconel 718 which had been given a similar
prerolling heat treatment followed by rolling reduction of 20 and 40% and
a post rolling heat treatment essentially as described in the Examples 2
and 3 above. The tensile properties of each of the samples are listed in
Table IV.
TABLE IV
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High Temperature Tensile Properties
Test Yield Tensile
Elonga-
Alloy Temp. Strength
Strength
tion
Test No.
%-CR (C) (ksi) (ksi) (%)
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1 CH83D 399 218.9 226.2 12.3
15% 593 207.1 220.4 9.6
704 190.1 190.0 24.8
2 CH83E 399 225.6 230.1 9.7
25% 593 220.0 230.9 5.9
704 198.4 206.1 27.2
3 CH83F 399 238.1 243.0 6.8
35% 593 219.8 229.7 7.6
704 205.3 212.2 24.8
4 CH84D 399 163.3 180.9 20.0
15% 593 145.9 166.5 14.3
704 138.9 154.2 34.3
5 CH84E 399 171.5 186.2 15.9
25% 593 163.3 186.3 14.6
704 157.0 165.9 33.5
6 CH84F 399 182.4 194.1 13.8
35% 593 164.5 180.6 12.2
704 155.3 165.2 33.4
7 IN718 649 187.8 195.9 10.8
20% 704 169.4 177.6 18.4
8 IN718 649 193.7 201.9 10.8
40% 704 187.2 194.9 25.0
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Referring now to Table IV the strength of the alloys IN718, CH84 and CH83
are compared.
Comparison is based initially on the comparison of the results of tests 2,
5 and 7. The reason for this comparison is that the degree of cold rolled
reduction in thickness if comparable for these three tests. Test 2
involved testing after a 25% reduction of alloy CH83. Test 5 involved
testing after a 25% reduction of alloy 84 and test 7 involved testing
after a 20% reduction of alloy IN718.
At 704.degree. C. the yield strength found for alloy IN718 of test 7 is
substantially stronger than the CH84E alloy of test 5 by about 12 ksi.
However the 704.degree. C. yield strength of the alloy 83E is very
surprisingly higher than that of the alloy 718 of test 7 and is in fact
about 30 ksi higher.
The significance of a 30 ksi gain in yield strength can be appreciated that
this represents about the total yield strength of conventional stainless
steels.
The 704.degree. C. tensile strength of the same alloys follows the same
pattern with the CH84B alloy showing substantially lower tensile strength
(about 10 ksi) than the IN718 and with the CH83B alloy of test 2
displaying a surprisingly greater tensile strength than the comparable
IN718 alloy sample of test 7.
In essentially all tests made the CH83 alloy displayed substantially higher
strength than the IN718 alloy while at the same time displaying fully
adequate ductility.
From the results listed in Table IV, it is clear that the alloy CH83, which
contains tantalum as a hardening element, shows excellent tensile
strengths up to about 704.degree. C. In contrast to the excellent tensile
properties of the CH83 alloy, the CH84 alloy which contained no tantalum
has much poorer tensile properties and is much weaker than the CH83 alloy.
Further, it can be observed from the results listed in Table IV that the
CH84 alloy which contains no tantalum is weaker than the Inconel 718 even
though the CH84 has about the same level of hardening elements. Hardening
element additions are commonly known, and known from the Eiselstein patent
U.S. Pat. No. 3,046,108 to be aluminum, titanium and niobium.
Further tests results were obtained for the alloys. In particular, stress
rupture results were obtained by conventional stress rupture measurements
and the results are plotted in FIG. 8.
The new alloys CH83 and CH84 exhibit the obvious advantage of temperature
capability over Inconel 718. The alloy CH83 with the tantalum additions
has an approximately 100.degree. F. temperature capability improvement
over that of the Inconel 718 alloy.
With further reference to FIG. 8 the IN718 alloy rupture life is seen to
increase slightly for the alloy cold rolled 40% over the alloy cold rolled
20%. As the inverted triangle (for 40%CR) stands above the upright
triangle (for 20%CR). The +, x and * data points for the CH84 alloy are
substantially above the triangles of the IN718 alloy. The square, diamond
and octagon data points for the CH83 alloy are substantially above the
CH84 data points and are quite far above the IN718 triangle data points.
This and other rupture life data confirm that the CH83 alloy has a
100.degree. F. temperature advantage over the IN718 alloy.
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